EP2511612B1 - Gas turbine component assembly - Google Patents

Gas turbine component assembly Download PDF

Info

Publication number
EP2511612B1
EP2511612B1 EP10835729.4A EP10835729A EP2511612B1 EP 2511612 B1 EP2511612 B1 EP 2511612B1 EP 10835729 A EP10835729 A EP 10835729A EP 2511612 B1 EP2511612 B1 EP 2511612B1
Authority
EP
European Patent Office
Prior art keywords
stator vanes
sidewalls
turbine stator
stage turbine
turbine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
EP10835729.4A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP2511612A1 (en
EP2511612A4 (en
Inventor
Rosic Budmir
Yasuro Sakamoto
Sumiu Uchida
Eisaku Ito
Tsuyoshi Kitamura
Satoshi Hada
Sosuke Nakamura
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Mitsubishi Power Ltd
Original Assignee
Mitsubishi Hitachi Power Systems Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Mitsubishi Hitachi Power Systems Ltd filed Critical Mitsubishi Hitachi Power Systems Ltd
Publication of EP2511612A1 publication Critical patent/EP2511612A1/en
Publication of EP2511612A4 publication Critical patent/EP2511612A4/en
Application granted granted Critical
Publication of EP2511612B1 publication Critical patent/EP2511612B1/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/60—Support structures; Attaching or mounting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/346—Feeding into different combustion zones for staged combustion

Definitions

  • the present invention relates to a communicating structure between a combustor and a turbine portion and to a gas turbine.
  • a gas turbine generally includes a compressor, a combustor, and a turbine portion as main components; the compressor is coupled to a turbine with a rotating shaft; and the combustor is disposed between the compressor and the turbine portion.
  • air which is working fluid
  • the compressor which is rotationally driven by the rotating shaft, to be compressed therein, and the compressed air is introduced into the combustor.
  • Fuel is mixed with the compressed gas in the combustor, and high-temperature, high-pressure combustion gas is generated by combustion of the mixed air.
  • the combustion gas is expelled to the turbine portion from the combustor to rotationally drive the turbine portion.
  • the high-temperature working fluid expelled from the combustor which includes the combustion gas, passes through between first-stage turbine stator vanes in the turbine portion and subsequently flows to first-stage turbine rotor blades.
  • first-stage turbine rotor blades part of the energy possessed by the working fluid is converted to rotational energy and is transmitted to the rotating shaft as a rotational driving force.
  • a rear end of a tail pipe of the combustor and leading edges of the first-stage turbine stator vanes positioned most upstream of the turbine portion are disposed with gaps therebetween. Accordingly, there is a problem in that part of the high-temperature working fluid that flows toward the turbine portion from the combustor flows into the gaps between the rear end of the tail pipe and the leading edges of the first-stage turbine stator vanes, and a loss occurs caused by this flow.
  • Patent Literature 1 Although integration of the combustor and the first-stage turbine stator vanes is described, there is no disclosure about the shape of an inner wall of the combustor inside which the high-temperature working fluid flows.
  • the present invention has been conceived in order to solve the above-described problems, and an object thereof is to provide a gas turbine component assembly according to claim 1 that is capable of suppressing the occurrence of a loss and also capable of reducing the flow level of cooling fluid employed in cooling turbine blades.
  • a communicating structure between combustors and a turbine portion according to the gas turbine component assembly as claimed in claim 1 is a communicating structure between combustors that generate combustion gas by combusting compressed air supplied from a compressor and fuel supplied from fuel nozzles, which are mixed inside a plurality of pipe pieces disposed next to each other around a rotating shaft, and a turbine portion that generates a rotational driving force by making the combustion gas sequentially pass through a turbine stage formed of a plurality of turbine stator vanes and turbine rotor blades disposed around the rotating shaft, wherein at least some of first-stage turbine stator vanes closest to the combustors among the turbine stator vanes are disposed downstream of sidewalls of one pipe piece and another pipe piece that are adjacent to each other, and the distance from leading edges of the first-stage turbine stator vanes disposed downstream of the sidewalls to end portions of the sidewalls closer to the turbine portion is equal to or less than
  • the communicating structure between the combustors and the turbine portion With the communicating structure between the combustors and the turbine portion, by disposing the first-stage turbine stator vanes positioned downstream of the sidewalls close to the end portions of the sidewalls closer to the turbine portion, the combustion gas is prevented from flowing in between the sidewalls and the first-stage turbine stator vanes. Accordingly, the occurrence of a loss due to the inflow of the combustion gas between the sidewalls and the first-stage turbine stator vanes is suppressed.
  • the leading edges of the first-stage turbine stator vanes are disposed in relatively cool flows behind (in the wake of) the sidewalls, and thus, the high-temperature combustion gas is less likely to directly collide with the leading edges of the first-stage turbine stator vanes. Accordingly, the need to cool the leading edges of the first-stage turbine stator vanes is reduced, and the flow level of the cooling fluid required for cooling is reduced.
  • the internal surfaces of the sidewalls have shapes that are smoothly continuous with external surfaces of the first-stage turbine stator vanes disposed downstream of the sidewalls.
  • the combustion gas generated inside the pipe pieces flows along the internal surfaces of the sidewalls and subsequently flows along the external surfaces of the first-stage turbine stator vanes that are smoothly continuous with the sidewalls. Accordingly, as compared with the case in which level differences, etc. are formed between the internal surfaces of the sidewalls and the external surfaces of the first-stage turbine stator vanes thereby making them discontinuous, the flow of combustion gas is less likely to be disturbed and the occurrence of the loss can be suppressed.
  • the flow of combustion gas at the external surfaces of the first-stage turbine stator vanes is less likely to be disturbed, for example, in the method in which the first-stage turbine stator vanes are cooled by making the cooling fluid flow in the form of a film at the external surfaces of the first-stage turbine stator vanes, deterioration in the efficiency of cooling the first-stage turbine stator vanes can be suppressed.
  • leading edges of the first-stage turbine stator vanes where the temperature thereof most easily reaches a high temperature, are protected by the sidewalls (disposed in the wake of the side walls), exposure to the high-temperature combustion gas is prevented, and the flow level of the cooling fluid required to cool the first-stage turbine stator vanes can be reduced.
  • the number of cooling holes from which cooling fluid employed to cool the first-stage turbine stator vanes is made to flow out to the peripheries of the first-stage turbine stator vanes be smaller in the first-stage turbine stator vanes disposed downstream of the sidewalls.
  • the combustion gas is less likely to collide with the leading edges of the first-stage turbine stator vanes disposed downstream of the sidewalls, as compared with the first-stage turbine stator vanes disposed elsewhere. Accordingly, as compared with the first-stage turbine stator vanes disposed at locations other than downstream of the sidewalls, it is possible to reduce the number of cooling holes or shower-head cooling holes, from which the cooling fluid is made to flow out to the peripheries of the first-stage turbine stator vanes disposed downstream of the sidewalls so as to flow along the external surfaces of the first-stage turbine stator vanes in the form of a film. In other words, as compared with the first-stage turbine stator vanes disposed at locations other than downstream of the sidewalls, the flow level of the cooling fluid employed to cool the first-stage turbine stator vanes can be reduced.
  • the cooling fluid for cooling the sidewalls be made to flow through a gap between a sidewall of the one pipe piece and a sidewall of the other pipe piece and that the cooling fluid that has cooled the sidewalls subsequently flow along the peripheries of the first-stage turbine stator vanes disposed downstream of the sidewalls from downstream-side end portions of the sidewalls.
  • the cooling fluid that has flowed between the sidewalls and cooled the sidewalls flow along the peripheries of the first-stage turbine stator vanes in the form of a film from the outflow channels, which are slot-like gaps formed between the downstream-side end portions of the sidewalls and the first-stage turbine stator vanes, the first-stage turbine stator vanes disposed downstream of the sidewalls can be effectively cooled by the cooling fluid. Accordingly, the flow level of the cooling fluid that is supplied to the first-stage turbine stator vanes disposed downstream of the sidewalls and that cools the first-stage turbine stator vanes can be reduced.
  • downstream-side end portions of the sidewalls be tilted in the direction in which the combustion gas is deflected by the first-stage turbine stator vanes.
  • the flow of combustion gas can be deflected by the downstream-side end portions of the sidewalls and the first-stage turbine stator vanes.
  • the size of the communicating structure between the combustors and the turbine portion in the axial direction of the rotating shaft can be reduced.
  • the deflection by the sidewalls can be increased, the deflection by the first-stage turbine stator vanes can be reduced; therefore, the axial-direction size can be further reduced.
  • the tilted portions of the sidewalls in cross-sectional view, form airfoil shapes together with the first-stage turbine stator vanes disposed downstream of the sidewalls.
  • the tilted portions of the sidewalls have cross-sectional shapes that form airfoil shapes together with the first-state turbine stator vanes, the flow of the combustion can be effectively deflected as compared with the case in which the airfoil shapes are not formed.
  • a gas turbine as claimed in claim 8 according to the present invention is a gas turbine including a compressor that compresses air; a gas turbine component assembly according to the invention that generates combustion gas by combusting compressed air supplied from the compressor and fuel supplied from a fuel nozzle, which are mixed therein; and a rotating shaft that transmits the rotational driving force from the turbine portion to the compressor.
  • the gas turbine according to the present invention because it has the communicating portion between the combustors and the turbine portion according to the present invention described above, the occurrence of a loss can be suppressed and the flow level of the cooling volume employed to cool the turbine stator vanes can be reduced; therefore, the efficiency of the gas turbine as a whole can be improved.
  • an advantage is afforded in that, by disposing first-stage turbine stator vanes positioned downstream of sidewalls closer to end portions of the sidewalls close to a turbine-portion, the occurrence of a loss in a gas turbine can be suppressed, and the flow level of cooling fluid employed to cool turbine blades can also be reduced.
  • Fig. 1 is a schematic view for explaining the configuration of a gas turbine according to this embodiment.
  • a gas turbine 1 of the present invention will be described as applied to one that drives a generator G.
  • the object to be driven by the gas turbine 1 is not limited to the generator G, and it may be other equipment; it is not particularly limited.
  • the gas turbine 1 is mainly provided with a compressor 2, combustors 3, a turbine portion 4, and a rotating shaft 5.
  • the compressor 2 takes in atmospheric air, which is external air, compresses the air, and supplies the compressed air to the combustors 3.
  • the compressor 2 is provided with an inlet guiding vane (not shown) which adjusts the flow level of the atmospheric air that flows into the compressor 2, first-stage rotor blades (not shown) that compress the atmospheric air that has flowed in, first-stage stator vanes (not shown), and so on.
  • Fig. 2 is a schematic view for explaining the configurations of the compressor, the turbine portion, and the combustor in Fig. 1 .
  • the combustors 3 are can-type combustors in which the air compressed by the compressor 2 and externally supplied fuel are mixed and that generates high-temperature combustion gas by combusting the mixed air that has been mixed therein.
  • the combustors 3 are mainly provided with air inlets 31, fuel nozzles 32, and tail pipes (tube pieces) 33.
  • the air inlets 31 guide the air compressed by the compressor 2 to the interior of the tail pipes 33 and are each disposed in the form of a ring around the fuel nozzles 32. Furthermore, the air inlets 31 impart the air flowing into the interior of the tail pipes 33 with flow speed components in a swirling direction, thus forming circulating flows inside the tail pipes 33.
  • the fuel nozzles 32 spray the externally supplied fuel into the interior of the tail pipes 33.
  • the fuel sprayed from the fuel nozzles 32 is stirred by the airflow formed by the air inlets 31, etc. to form mixed air containing fuel and air.
  • the tail pipes 33 are pipe-shaped members that extend toward an inflow portions of the turbine portion 4 from the air inlets 31 and the fuel nozzles 32.
  • the tail pipes 33 are where the mixed air containing fuel and air and the combustion gas generated by the combustion of the mixed air flow in the interior thereof.
  • the sectional shape of the tail pipes 33 near the fuel nozzles 32 is substantially circular, and the sectional shape thereof near the turbine portion 4 is substantially rectangular. Accordingly, the sectional shape of the tail pipes 33 continuously changes from a substantially circular shape to a substantially rectangular shape from the fuel nozzles 32 toward the turbine portion 4.
  • the turbine portion 4 receives a supply of high-temperature gas generated by the combustors 3 to generate a rotational driving force and transmits the generated rotational driving force to the rotating shaft 5.
  • Fig. 3 is a partially enlarged view for explaining the communicating structure between the combustor and the turbine portion in Fig. 1 .
  • the turbine portion 4 is provided with first-stage turbine stator vanes (turbine stator vanes) 4SV and first-stage turbine rotor blades (turbine rotor blades) 4RB.
  • the first-stage turbine stator vanes 4SV form a turbine stage together with the first-stage turbine rotor blades 4RB and generate the rotational driving force together with the first-stage rotor blades 4RB from the high-temperature gas that has flowed into the turbine portion 4.
  • the first-stage turbine stator vanes 4SV are a plurality of blades that are arranged around the rotating shaft at equal intervals at positions that face downstream-side end portions (bottom-side end portions in Fig. 3 ) of the tail pipes 33 with respect to a flow of combustion gas and that are also arranged so as to extend along a radial direction (vertical direction in Fig. 3 with respect to the plane of the drawing). Furthermore, the first-stage turbine stator vanes 4SV deflect the combustion gas that has flowed into a row of the first-stage turbine stator vanes 4SV from the combustors 3 in a circumferential direction (left-right direction in Fig. 3 ).
  • the number of first-stage turbine stator vanes 4SV is an integral multiple of the number of combustors 3, and, at least some of the first-stage turbine stator vanes 4SV are disposed downstream of sidewalls 34 of the tail pipes 33 in the combustors 3, as shown in Fig. 3 .
  • the first-stage turbine stator vanes 4SV are arranged so that a distance L from leading edges LE of the first-stage turbine stator vanes 4SV to end portions of the sidewalls 34 closer to the turbine portion 4 is set to be equal to or less than a thickness T which is the sum of the thicknesses of the sidewall 34 of one tail pipe 33 and the sidewall 34 of another tail pipe 33 that are adjacent to each other, and gaps between the two sidewalls 34 and 34 are combined, in other words, the thickness T (hereinafter, referred to as "thickness T related to the sidewalls 34") is the spacing between the inner surface of the sidewall 34 of one tail pipe 33 and the inner surface of the sidewall 34 of another tail pipe, which are adjacent to each other.
  • first-stage turbine stator vanes 4SV are provided with cavities 41 to which cooling air (cooling fluid) that protects the first-stage turbine stator vanes 4SV from the heat of the high-temperature gas flowing in the peripheries thereof is supplied and are provided with a plurality of cooling holes 42 that perform film cooling wherein the cooling air is made to flow out into the peripheries of the first-stage turbine stator vanes 4SV from the cavities 41.
  • cooling air cooling fluid
  • the cooling holes 42 are arranged in a large number at the leading edges LE of the first-stage turbine stator vanes 4SV, where the heat load is high, so that the leading edges LE are formed like shower heads.
  • the first-stage turbine rotor blades 4RB form the turbine stage together with the first-stage turbine stator vanes 4SV and generate the rotational driving force on the basis of the combustion gas deflected by the first-stage turbine stator vanes 4SV.
  • the first-stage turbine rotor blades 4RB are a plurality of blades that are arranged around the rotating shaft at equal intervals at positions downstream (right-side positions in Fig. 2 ) of the first-stage turbine stator vanes 4SV with respect to the flow of combustion gas and that are also arranged so as to extend along the radial direction (top-bottom direction in Fig. 2 ). Furthermore, the first-stage turbine rotor blades 4RB receive the combustion gas deflected by the first-stage turbine stator vanes 4SV and rotationally drive the rotating shaft 5.
  • cooling air that protects the first-stage turbine rotor blades 4RB from the heat of the combustion gas that flows in the peripheries thereof is supplied to the first-stage turbine rotor blades 4RB.
  • the turbine portion 4 may be provided only with the first-stage turbine stator vanes 4SV and the first-stage turbine rotor blades 4RB, as described above, or second-stage turbine stator vanes and second-stage turbine rotor blades, third-stage turbine stator vanes and third-stage turbine rotor blades, and so on may be additionally provided; it is not particularly limited.
  • the gas turbine 1 takes in the atmospheric air (air) when the compressor 2 is rotationally driven.
  • the air that has been taken in is compressed by the compressor 2 and is discharged toward the combustor 3.
  • the compressed air that has flowed into the combustors 3 is mixed with the externally supplied fuel at the combustors 3.
  • the mixed air containing fuel and air is combusted in the combustors 3, and the combustion gas is generated.
  • the combustion gas generated in the combustors 3 is supplied to the turbine portion 4 downstream of the combustors 3.
  • the combustion gas flows out from the tail pipes 33 of the combustors 3 and flows into the row of the first-stage turbine stator vanes 4SV in the turbine 4.
  • the combustion gas is less likely to flow in between the first-stage turbine stator vanes 4SV disposed downstream of the sidewalls 34 of the tail pipes 33 and the tail pipes 33, and a loss due to this flow is less likely to occur.
  • leading edges LE of the first-stage turbine stator vanes 4SV disposed downstream of the sidewalls 34 are positioned in flows behind (in the wake of) the sidewalls 34; therefore, the combustion gas is less likely to directly collide with the leading edges LE.
  • the flow direction of the combustion gas that has flowed into the row of the first-stage turbine stator vanes 4SV is deflected in the circumferential direction (left-right direction in Fig. 3 ), centered around the rotating shaft 5, and flows into the row of the first-stage turbine rotor blades 4RB, as shown in Fig. 2 .
  • the first-stage turbine rotor blades 4RB are rotationally driven by the deflected combustion gas.
  • the rotational driving force generated in this way at the turbine portion 4 is transmitted to the rotating shaft 5.
  • the rotating shaft 5 transmits the rotational driving force extracted at the turbine portion 4 to the compressor 2 and the generator G.
  • the first-stage turbine stator vanes 4SV positioned downstream of the sidewalls 34 are disposed close to the end portions of the sidewalls 34 closer to the turbine portion 4, and thereby, the combustion gas is prevented from flowing in between the sidewalls 34 and the first-stage turbine stator vanes 4SV. Because of this, the occurrence of loss caused by having the combustion gas flow in between the sidewalls 34 and the first-stage turbine stator vanes 4SV can be suppressed.
  • the leading edges LE of the first-stage turbine stator vanes 4SV are disposed in relatively cool flows behind (in the wake of) the sidewalls 34, and the high-temperature combustion gas is less likely to directly collide with the leading edges LE of the first-stage turbine stator vanes 4SV. Because of this, the need to cool the leading edges LE of the first-stage turbine stator vanes 4SV is reduced, and the flow level of the cooling air required for cooling can be reduced.
  • the combustion gas is less likely to collide with the first-stage turbine stator vanes 4SV disposed downstream of the sidewalls 34 at the leading edges LE thereof as compared with the first-stage turbine stator vanes 4SV disposed elsewhere. Accordingly, as compared with the first-stage turbine stator vanes 4SV disposed at locations other than downstream of the sidewalls 34, it is possible to reduce the number of cooling holes 42 in the first-stage turbine stator vanes 4SV, which cause the cooling air to flow along external surfaces thereof in the form of a film by making the cooling air flow out therefrom to the peripheries of the first-stage turbine stator vanes 4SV.
  • Fig. 4 is a partially enlarged view for explaining the communicating structure between the combustors and the turbine portion in the gas turbine according to this embodiment.
  • combustors 103 in a gas turbine 101 in this embodiment differ from those of the first embodiment in the shapes of the end portions (bottom-side end portions in Fig. 4 ) of sidewalls 134 of tail pipes (pipe pieces) 133 closer to a turbine portion 104.
  • cooling channels 145 in which cooling fluid (for example, compressed air compressed by the compressor 2), such as cooling air, etc., flows and that extend in a direction (top-bottom direction in Fig. 4 ) in which the combustion gas flows are provided between the tail pipes 133 of adjacent combustors 103.
  • cooling fluid for example, compressed air compressed by the compressor 2
  • cooling air such as cooling air, etc.
  • End portions of the cooling channels 145 closer to the turbine portion 104 are opened at the end portions (bottom-side end portions in Fig. 4 ) of the sidewalls 134 of the end pipes 133 closer to the turbine portion 104.
  • Fig. 5 is an enlarged view for explaining the configurations of the sidewalls and the first-stage turbine stator vanes in Fig. 4 .
  • the downstream-side end portions of the sidewalls 134 are formed in shapes such that internal surfaces of the sidewalls 134 are smoothly continuous with external surfaces of first-stage turbine stator vanes 104SV adjacent thereto.
  • the sidewalls 134 are formed so that the widths of the sidewalls 134 are increased toward the first-stage turbine stator vanes 104SV.
  • first-stage turbine stator vanes 4SV and first-stage turbine stator vanes (turbine stator vanes) 104SV are provided at the turbine portion 104 in the gas turbine 101 of this embodiment, as shown in Fig. 4 .
  • the first-stage turbine stator vanes 4SV and the first-stage turbine stator vanes 104SV form a turbine stage together with the first-stage turbine rotor blades 4RB and generate a rotational driving force together with the first-stage rotor blades 4RB from the combustion gas that has flowed into the turbine portion 104. Furthermore, the first-stage turbine stator vanes 4SV and the first-stage turbine stator vanes 104SV are a plurality of blades that are arranged at equal intervals on the same circumference around the rotating shaft 5 and that are also arranged so as to extend along the radial direction (vertical direction in Fig. 4 with respect to the plane of the drawing).
  • the first-stage turbine stator vanes 4SV are turbine stator vanes disposed between the sidewalls 134, in other words, turbine stator vanes disposed between the first-stage turbine stator vanes 104SV.
  • the first-stage turbine stator vanes 104SV are turbine stator vanes disposed at positions facing the downstream-side end portions (bottom-side end portions in Fig. 4 ) of the end pipes 133 with respect to the flow of combustion gas, in other words, turbine stator vanes disposed between the first-stage turbine stator vanes 4SV.
  • the cavities 41 inside which the cooling air is supplied and the cooling holes 42 from which the cooling air from the cavities 41 is made to flow out to the peripheries of the first-stage turbine stator vanes 104SV are not formed in the first-stage turbine stator vanes 104SV.
  • outflow channels 146 that communicate with the cooling channels 145 at the sidewalls 134 and from which the cooling air, after flowing through the cooling channels 145, flows out along the peripheries of the first-stage turbine stator vanes 104SV in the form of a film are provided between the first-stage turbine stator vanes 104SV and the sidewalls 134.
  • the outflow channels 146 are long, narrow slots that extend from the cooling channels 145 toward the outer side of the sidewalls 134 in the downstream direction (right direction in Fig. 5 ) of the flow of combustion gas.
  • the combustion gas flows out from the tail pipes 133 of the combustors 103 and flows into a row of the first-stage turbine stator vanes 4SV and the first-stage turbine stator vanes 104SV at the turbine portion 104.
  • the combustion gas that has flowed along the internal surfaces of the sidewalls 134 of the tail pipes 133 is deflected while flowing along the external surfaces of first-stage turbine stator vanes 104SV from the internal surface of the sidewalls 134.
  • the cooling air that has flowed through the cooling channels 145 and cooled the tail pipes 133 flows out along the external surfaces of the first-stage turbine stator vanes 104SV via the outflow channels 146.
  • the cooling air flows along the external surfaces of the first-stage turbine stator vanes 104SV in the form of a film and cools the first-stage turbine stator vanes 104SV.
  • the combustion gas that has flowed through the centers of the tail pipes 133 collides with the first-stage turbine stator vanes 4SV and is deflected while flowing along the surfaces of the first-stage turbine stator vanes 4SV.
  • the combustion gas generated inside the tail pipes 133 flows along the internal surfaces of the sidewalls 134 and subsequently flows along the external surfaces of the first-stage turbine stator vanes 104SV, which are smoothly continuous therewith. Accordingly, as compared with the case in which the internal surfaces of the sidewalls 134 and the external surfaces of the first-stage turbine stator vanes 104SV are discontinuous due to the formation of a level difference, etc. therebetween, the flow of combustion gas is less likely to be disturbed, and loss can be suppressed.
  • the flow of combustion gas at the external surfaces of the first-stage turbine stator vanes 104SV is less likely to be disturbed, with the approach in which the cooling air that has flowed out from the outflow channels 146 is made to flow at the external surfaces of the first-stage turbine stator vanes 104SV in the form of a film to cool the first-stage turbine stator vanes 104SV, deterioration of the cooling efficiency of the first-stage turbine stator vanes 104SV can be prevented.
  • the first-stage turbine stator vanes 104SV disposed downstream of the sidewalls 134 can be cooled with the cooling air. Accordingly, it is possible to reduce the flow level of the cooling air to be supplied to the first-stage turbine stator vanes 104SV to cool the first-stage turbine stator vanes 104SV.
  • Fig. 6 is a partially enlarged view for explaining the communicating structure between the combustors and the turbine portion in the gas turbine according to this embodiment.
  • a combustor 203 in a gas turbine 201 of this embodiment differs from that in the first embodiment in the shapes of the end portions (bottom-side end portions in Fig. 6 ) of sidewalls 234 of tail pipes (pipe pieces) 233 closer to a turbine portion 204.
  • the sidewalls 234 of the tail pipes 233 in the combustors 203 are provided with tilted portions 235 that are tilted in the direction in which the first-stage turbine stator vanes 4SV deflect the flow of combustion gas.
  • the tilted portions 235 are end portions of the sidewalls 234 closer to the turbine portion 204 and are portions adjacent to the first-stage turbine stator vanes 204SV. Furthermore, because the tilted portions 235 are formed by tilting the sidewalls 234 without other modifications, the thickness-wise size of the tilted portions 235 and the thickness-wise size of parts of the sidewalls 234 other than the tilted portions 235 are the same.
  • the tail pipes 233 and the sidewalls 234 are provided with cooling channels 145 that extend along the direction in which the combustion gas flows (top-bottom direction in Fig. 6 ) and inside which cooling fluid, such as cooling air, etc., flows. Furthermore, the cooling channels 145 extend along the tilted portions 235 inside the tilted portions 235 of the sidewalls 234.
  • the end portions of the cooling channels 145 closer to the turbine portion 204 open at the end portions (bottom-side end portions in Fig. 6 ) of the tilted portions 235 of the sidewalls 234 closer to the turbine portion 204.
  • the turbine portion 204 of the gas turbine 201 in this embodiment is provided with the first-stage turbine stator vanes 4SV and the first-stage turbine stator vanes (turbine stator vanes) 204SV.
  • the first-stage turbine stator vanes 4SV and the first-stage turbine stator vanes 204SV form a turbine stage together with the first-stage turbine rotor blades 4RB and generate a rotational driving force together with the first-stage rotor blades 4RB from the combustion gas that has flowed into the turbine portion 204. Furthermore, the first-stage turbine stator vanes 4SV and the first-stage turbine stator vanes 204SV are a plurality of blades that are arranged at equal intervals on the same circumference around the rotating shaft 5 and that are also arranged so as to extend along the radial direction (vertical direction in Fig. 6 with respect to the plane of the drawing).
  • the first-stage turbine stator vanes 4SV are turbine stator vanes disposed between the sidewalls 234 and the tilted portions 235, in other words, turbine stator vanes disposed between the first-stage turbine stator vanes 204SV.
  • the first-stage turbine stator vanes 204SV are turbine stator vanes disposed at positions facing the downstream-side end portions (bottom-side end portions in Fig. 6 ) of the tilted portions 235 with respect to the flow of combustion gas, in other words, turbine stator vanes disposed between the first-stage turbine stator vanes 4SV.
  • the first-stage turbine stator vanes 204SV are formed with a smaller sectional area as compared with the first-stage turbine stator vanes 4SV, and a portion in the first-stage turbine stator vanes 204SV where the thickness-wise size is the largest has the same thickness-wise size as the tilted portions 235.
  • the cavities 41 inside which the cooling air is supplied and the cooling holes 42 from which the cooling air from the cavities 41 is made to flow out to the peripheries of the first-stage turbine stator vanes 204SV are not formed in the first-stage turbine stator vanes 204SV.
  • outflow channels 146 that communicate with the cooling channels 145 at the tilted portions 235 and from which the cooling air, after flowing through the cooling channels 145, flows out along the peripheries of the first-stage turbine stator vanes 204SV are provided between the first-stage turbine stator vanes 204SV and the tilted portions 235.
  • the outflow channels 146 are through-holes that extend from the cooling channels 145 toward the outer side of the tilted portions 235 in the downstream direction (left-bottom direction in Fig. 6 ) of the flow of combustion gas.
  • the combustion gas flows out from the tail pipes 233 of the combustors 203 and flows into the row of the first-stage turbine stator vanes 4SV and the first-stage turbine stator vanes 204SV at the turbine portion 204.
  • the combustion gas that has flowed along the internal surfaces of the sidewalls 234 of the tail pipes 233 is deflected while flowing along the internal surfaces of the tilted portions 235 at the sidewalls 234 and the external surfaces of first-stage turbine stator vanes 204SV.
  • the cooling air that has flowed through the cooling channels 145 and cooled the tail pipes 233 and the tilted portions 235 flows out along the external surfaces of the first-stage turbine stator vanes 204SV via the outflow channels 146.
  • the cooling air flows along the external surfaces of the first-stage turbine stator vanes 204SV in the form of a film and cools the first-stage turbine stator vanes 204SV.
  • the combustion gas that has flowed through the interior of the tail pipes 233 collides with the first-stage turbine stator vanes 4SV and is deflected while flowing along the surfaces of the first-stage turbine stator vanes 4SV.
  • the flow of combustion gas can be deflected by the tilted portions 235, which are the downstream-side end portions of the sidewalls 234, and the first-stage turbine stator vanes 204SV.
  • the deflection by the sidewalls 234 can be further increased, the deflection by the first-stage turbine stator vanes 204SV can be reduced; therefore, the size in the axial direction of the rotating shaft 5 can be further reduced.
  • Fig. 7 is a partially enlarged view for explaining the communicating structure between the combustors and the turbine portion in the gas turbine according to this embodiment.
  • a turbine portion 304 in a gas turbine 101 of this embodiment differs from that in the first embodiment in the shapes and arrangement of first-stage turbine stator vanes (turbine stator vanes) 304SV.
  • the first-stage turbine stator vanes 304SV form a turbine stage together with the first-stage turbine rotor blades 4RB and generate a rotational driving force together with the first-stage rotor blades 4RB from the combustion gas that has flowed into the turbine portion 304. Furthermore, the first-stage turbine stator vanes 304SV are a plurality of blades that are arranged at equal intervals on the same circumference around the rotating shaft 5 and that are also arranged so as to extend along the radial direction (vertical direction in Fig. 7 with respect to the plane of the drawing).
  • the first-stage turbine stator vanes 304SV are disposed at positions facing the downstream-side end portions (bottom-side end portions in Fig. 7 ) of the sidewalls 334 of tail pipes 333 with respect to the flow of combustion gas. In other words, the first-stage turbine stator vanes 304SV are provided in the same number as the number of combustors 303.
  • the first-stage turbine stator vanes 304SV have similar shapes to the first-stage turbine stator vanes 4SV in the first embodiment, etc. and are formed with larger sectional areas.
  • leading edges LE of the first-stage turbine stator vanes 304SV are disposed at positions separated from the downstream-side end portions of the sidewalls 334, at most, by the thickness T related to the sidewalls 334, and trailing edges TE of the first-stage turbine stator vanes 304SV are disposed at the same positions as trailing edges TE of conventional first-stage turbine stator vanes.
  • the combustion gas flows out from the tail pipes 333 of the combustors 103 and flows into the row of the first-stage turbine stator vanes 304SV at the turbine portion 304.
  • the combustion gas that has flowed along the internal surfaces of the sidewalls 334 of the tail pipes 333 is deflected while flowing along the external surfaces of the first-stage turbine stator vanes 304SV.
  • the cooling air that has flowed through the cooling channels 145 and cooled the tail pipe 333 flows out along the external surfaces of the first-stage turbine stator vanes 304SV from the downstream-side end portions of the sidewalls 334.
  • the cooling air flows along the external surfaces of the first-stage turbine stator vanes 304SV in the form of a film and cools the first-stage turbine stator vanes 304SV.
  • the number of the first-stage turbine stator vanes 304SV can be reduced. Accordingly, a reduction in flow speed of the combustion gas due to friction or the like that acts between the first-stage turbine stator vanes 304SV and the combustion gas can be suppressed, and the loss caused by this can be suppressed.
  • Fig. 8 is a partially enlarged view for explaining the communicating structure between the combustors and the turbine portion in the gas turbine according to this embodiment.
  • combustors 403 in a gas turbine 401 of this embodiment differ from those in the first embodiment in the shapes of the end portions (bottom-side end portions in Fig. 8 ) of sidewalls 434 of tail pipes (pipe pieces) 433 closer to a turbine portion 404.
  • the sidewalls 434 of the tail pipes 433 in the combustors 403 are provided with tilted portions 435 that deflect the flow of combustion gas leftward in Fig. 8 .
  • the tilted portions 435 are end portions of the sidewalls 434 closer to the turbine portion 404 and are portions adjacent to the first-stage turbine stator vanes 404SV. Furthermore, the tilted portions 435 are formed in shapes whose cross-sections form airfoil shapes together with the first-stage turbine stator vanes 404SV.
  • upstream-side end portions (top-side end portions in Fig. 8 ) of the tilted portions 435 with respect to the flow of combustion gas are at positions equivalent to the leading edges LE of the first-stage turbine stator vanes 304SV in the fourth embodiment.
  • the cooling channels 145 in which cooling fluid (for example, compressed air compressed in the compressor 2), such as cooling air, flows and that extend along the direction (top-down direction in Fig. 8 ) in which the combustion gas flows are provided between adjacent tail pipes 433. Furthermore, the cooling channels 145 extend along the tilted portions 435, between the tilted portions 435 of adjacent sidewalls 434.
  • cooling fluid for example, compressed air compressed in the compressor 2
  • the cooling channels 145 extend along the tilted portions 435, between the tilted portions 435 of adjacent sidewalls 434.
  • End portions of the cooling channels 145 open at downstream-side end portions (bottom-side end portions in Fig. 8 ) of the tilted portions 435 of the sidewalls 434.
  • the turbine portion 404 of the gas turbine 401 in this embodiment is provided with the first-stage turbine stator vanes (turbine stator vanes) 404SV.
  • the first-stage turbine stator vanes 404SV form a turbine stage together with the first-stage turbine rotor blades 4RB and generate a rotational driving force together with the first-stage rotor blades 4RB from the combustion gas that has flowed into the turbine portion 404. Furthermore, the first-stage turbine stator vanes 404SV are a plurality of blades that are arranged at equal intervals on the same circumference around the rotating shaft 5 and that are also arranged so as to extend along the radial direction (vertical direction in Fig. 8 with respect to the plane of the drawing).
  • the first-stage turbine stator vanes 404SV are turbine stator vanes disposed at positions facing the downstream-side end portions (bottom-side end portions in Fig. 8 ) of the tilted portions 435 with respect to the flow of combustion gas.
  • the first-stage turbine stator vanes 404SV are formed with smaller sectional areas as compared with the first-stage turbine stator vanes 4SV in the first embodiment and form airfoil shapes together with the tilted portions 435.
  • trailing edges TE of the first-stage turbine stator vanes 404SV are disposed at the same positions as the trailing edges TE of the first-stage turbine stator vanes 4SV in the first embodiment, etc.
  • the cavities 41 inside which the cooling air is supplied and the cooling holes 42 from which the cooling air from the cavities 41 is made to flow out to the peripheries of the first-stage turbine stator vanes 404SV are not formed in the first-stage turbine stator vanes 404SV.
  • the outflow channels 146 that communicate with the cooling channels 145 and from which the cooling air, after flowing through the cooling channels 145, flows out along external surfaces of the first-stage turbine stator vanes 404SV in the form of a film are provided between the first-stage turbine stator vanes 404SV and the tilted portions 435.
  • the outflow channels 146 are long, narrow slots that extend from the cooling channels 145 toward the outer side of the tilted portions 435 in the downstream direction (left-bottom direction in Fig. 8 ) of the flow of combustion gas.
  • the combustion gas flows out from the tail pipes 433 of the combustors 403 and flows into the row of the first-stage turbine stator vanes 404SV at the turbine portion 404.
  • combustion gas that has flowed along the internal surfaces of the sidewalls 434 of the tail pipes 433 is deflected while flowing along the internal surfaces of the tilted portions 435 at the sidewalls 434 and the external surfaces of first-stage turbine stator vanes 404SV.
  • the cooling air that has flowed through the cooling channels 145 and cooled the tail pipes 433 and the tilted portions 435 flows out along the external surfaces of the first-stage turbine stator vanes 404SV via the outflow channels 146.
  • the cooling air flows along the external surfaces of the first-stage turbine stator vanes 404SV in the form of a film and cools the first-stage turbine stator vanes 404SV.
  • the cross-sections of the tilted portions 435 at the sidewalls 434 have shapes that form airfoil shapes together with the first-stage turbine stator vanes 404SV, as compared with the case in which the airfoil shapes are not formed, the flow of combustion gas can be effectively deflected.
  • Fig. 9 is a partially enlarged view for explaining the communicating structure between the combustors and the turbine portion in the gas turbine according to this embodiment.
  • combustors 503 in a gas turbine 501 of this embodiment differ from those in the first embodiment in the shapes of the end portions (bottom-side end portions in Fig. 9 ) of sidewalls 534 of tail pipes (pipe pieces) 533 closer to a turbine portion 504.
  • the sidewalls 534 of the tail pipes 533 in the combustors 503 are provided with tilted portions 535 that deflect the flow of combustion gas leftward in Fig. 9 .
  • the tilted portions 535 are end portions of the sidewalls 534 closer to the turbine portion 504 and are portions adjacent to the first-stage turbine stator vanes 504SV. Furthermore, because the tilted portions 535 are formed by tilting the sidewalls 534 without other modifications, the thickness-wise size of the tilted portions 535 and the thickness-wise size of parts of the sidewalls 534 other than the tilted portions 535 are the same.
  • upstream-side end portions (top-side end portions in Fig. 9 ) of the tilted portions 535 with respect to the flow of combustion gas are at positions equivalent to the leading edges LE of the first-stage turbine stator vanes 304SV in the fourth embodiment.
  • the cooling channels 145 inside which cooling fluid, such as the cooling air, flows and that extend in the direction (top-bottom direction in Fig. 9 ) in which the combustion gas flows are provided between adjacent tail pipes 533. Furthermore, the cooling channels 145 extend along the tilted portions 535, between the tilted portions 535 at the sidewalls 534.
  • End portions of the cooling channels 145 open at downstream-side end portions (bottom-side end portions in Fig. 9 ) of the tilted portions 535 at the sidewalls 534.
  • the turbine portion 504 of the gas turbine 501 in this embodiment is provided with the first-stage turbine stator vanes (turbine stator vanes) 504SV.
  • the first-stage turbine stator vanes 504SV form a turbine stage together with the first-stage turbine rotor blades 4RB and generate a rotational driving force together with the first-stage rotor blades 4RB from the combustion gas that has flowed into the turbine portion 504. Furthermore, the first-stage turbine stator vanes 504SV are a plurality of blades that are arranged at equal intervals on the same circumference around the rotating shaft 5 and that are also arranged so as to extend along the radial direction (vertical direction in Fig. 9 with respect to the plane of the drawing).
  • the first-stage turbine stator vanes 504SV are disposed at positions facing the downstream-side end portions (bottom-side end portions in Fig. 9 ) of the tilted portions 535 with respect to the flow of combustion gas.
  • the first-stage turbine stator vanes 504SV are formed with a smaller sectional area as compared with the first-stage turbine stator vanes 4SV, and a portion in the first-stage turbine stator vanes 504SV where the thickness-wise size is the largest has the same thickness-wise size as the tilted portions 535.
  • trailing edges TE of the first-stage turbine stator vanes 504SV are disposed at the same positions as the trailing edges TE of the first-stage turbine stator vanes 4SV in the first embodiment, etc.
  • the cavities 41 inside which the cooling air is supplied and the cooling holes 42 from which the cooling air from the cavities 41 is made to flow out to the peripheries of the first-stage turbine stator vanes 504SV are not formed in the first-stage turbine stator vanes 504SV.
  • outflow channels 146 that communicate with the cooling channels 145 at the tilted portions 535 and from which the cooling air, after flowing through the cooling channels 145, flows out along the peripheries of the first-stage turbine stator vanes 504SV are provided between the first-stage turbine stator vanes 504SV and the tilted portions 535.
  • the outflow channels 146 are through-holes that extend from the cooling channels 145 toward the outer side of the tilted portions 535 in the downstream direction (left-bottom direction in Fig. 9 ) of the flow of combustion gas.
  • the combustion gas flows out from the tail pipes 533 of the combustors 503 and flows into the row of the first-stage turbine stator vanes 504SV at the turbine portion 504.
  • the combustion gas that has flowed along the internal surfaces of the sidewalls 534 of the tail pipes 533 is deflected while flowing along the internal surfaces of the tilted portions 535 at the sidewalls 534 and the external surfaces of first-stage turbine stator vanes 504SV.
  • the cooling air that has flowed through the cooling channels 145 and cooled the tail pipes 533 and the tilted portions 535 flows out along the external surfaces of the first-stage turbine stator vanes 504SV via the outflow channels 146.
  • the cooling air flows along the external surfaces of the first-stage turbine stator vanes 504SV in the form of a film and cools the first-stage turbine stator vanes 504SV.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP10835729.4A 2009-12-07 2010-05-14 Gas turbine component assembly Active EP2511612B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP2009277746A JP5479058B2 (ja) 2009-12-07 2009-12-07 燃焼器とタービン部との連通構造、および、ガスタービン
PCT/JP2010/058171 WO2011070806A1 (ja) 2009-12-07 2010-05-14 燃焼器とタービン部との連通構造、および、ガスタービン

Publications (3)

Publication Number Publication Date
EP2511612A1 EP2511612A1 (en) 2012-10-17
EP2511612A4 EP2511612A4 (en) 2016-05-11
EP2511612B1 true EP2511612B1 (en) 2017-11-08

Family

ID=44145362

Family Applications (1)

Application Number Title Priority Date Filing Date
EP10835729.4A Active EP2511612B1 (en) 2009-12-07 2010-05-14 Gas turbine component assembly

Country Status (6)

Country Link
US (1) US9395085B2 (zh)
EP (1) EP2511612B1 (zh)
JP (1) JP5479058B2 (zh)
KR (2) KR101377772B1 (zh)
CN (1) CN102686949B (zh)
WO (1) WO2011070806A1 (zh)

Families Citing this family (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP5848074B2 (ja) * 2011-09-16 2016-01-27 三菱日立パワーシステムズ株式会社 ガスタービン、尾筒及び燃焼器
JP5834876B2 (ja) * 2011-12-15 2015-12-24 株式会社Ihi インピンジ冷却機構、タービン翼及び燃焼器
JP5927893B2 (ja) * 2011-12-15 2016-06-01 株式会社Ihi インピンジ冷却機構、タービン翼及び燃焼器
JP5886040B2 (ja) * 2011-12-28 2016-03-16 三菱日立パワーシステムズ株式会社 ガスタービン
US20130236301A1 (en) * 2012-03-09 2013-09-12 General Electric Company Apparatus And System For Directing Hot Gas
US9133722B2 (en) * 2012-04-30 2015-09-15 General Electric Company Transition duct with late injection in turbine system
EP2725197A1 (en) 2012-10-24 2014-04-30 Alstom Technology Ltd Combustor transition
EP2725196A1 (en) 2012-10-24 2014-04-30 Alstom Technology Ltd Combustor transition
JP6194120B2 (ja) * 2014-08-19 2017-09-06 三菱日立パワーシステムズ株式会社 ガスタービン
EP3124749B1 (en) * 2015-07-28 2018-12-19 Ansaldo Energia Switzerland AG First stage turbine vane arrangement
US10641491B2 (en) * 2016-03-25 2020-05-05 General Electric Company Cooling of integrated combustor nozzle of segmented annular combustion system
US10641175B2 (en) * 2016-03-25 2020-05-05 General Electric Company Panel fuel injector
US11299994B2 (en) 2017-06-29 2022-04-12 Mitsubishi Power, Ltd. First-stage stator vane for gas turbine, gas turbine, stator vane unit for gas turbine, and combustor assembly
JP6934350B2 (ja) * 2017-08-03 2021-09-15 三菱パワー株式会社 ガスタービン
JP7154984B2 (ja) * 2018-12-11 2022-10-18 三菱重工業株式会社 タービン静翼及びガスタービン
JP7232035B2 (ja) * 2018-12-18 2023-03-02 三菱重工業株式会社 ガスタービンの静翼及びガスタービン
KR102096580B1 (ko) 2019-04-01 2020-04-03 두산중공업 주식회사 예혼합 균일성이 향상된 연소기 노즐 및 이를 구비하는 가스터빈용 연소기
US11371702B2 (en) 2020-08-31 2022-06-28 General Electric Company Impingement panel for a turbomachine
US11614233B2 (en) 2020-08-31 2023-03-28 General Electric Company Impingement panel support structure and method of manufacture
US11460191B2 (en) 2020-08-31 2022-10-04 General Electric Company Cooling insert for a turbomachine
US11994292B2 (en) 2020-08-31 2024-05-28 General Electric Company Impingement cooling apparatus for turbomachine
US11994293B2 (en) 2020-08-31 2024-05-28 General Electric Company Impingement cooling apparatus support structure and method of manufacture
US11255545B1 (en) 2020-10-26 2022-02-22 General Electric Company Integrated combustion nozzle having a unified head end
US11767766B1 (en) 2022-07-29 2023-09-26 General Electric Company Turbomachine airfoil having impingement cooling passages

Family Cites Families (40)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2743579A (en) * 1950-11-02 1956-05-01 Gen Motors Corp Gas turbine engine with turbine nozzle cooled by combustion chamber jacket air
US3018624A (en) * 1954-03-02 1962-01-30 Bristol Siddeley Engines Ltd Flame tubes for use in combustion systems of gas turbine engines
GB980363A (en) * 1961-12-04 1965-01-13 Jan Jerie Improvements in or relating to gas turbines
JPS501646B1 (zh) * 1970-07-11 1975-01-20
US4852355A (en) * 1980-12-22 1989-08-01 General Electric Company Dispensing arrangement for pressurized air
JPS616606U (ja) * 1984-06-19 1986-01-16 三菱重工業株式会社 ガスタ−ビン燃焼器の翼冷却機構
JPS616606A (ja) * 1984-06-21 1986-01-13 Matsushita Electric Ind Co Ltd 光導波路の製造方法
US4739621A (en) 1984-10-11 1988-04-26 United Technologies Corporation Cooling scheme for combustor vane interface
JPS62121835A (ja) * 1985-11-21 1987-06-03 Agency Of Ind Science & Technol 高温空冷ガスタ−ビン
JPH0617652B2 (ja) 1988-02-05 1994-03-09 株式会社日立製作所 ガスタービン燃焼器
US5458461A (en) * 1994-12-12 1995-10-17 General Electric Company Film cooled slotted wall
JP3621216B2 (ja) * 1996-12-05 2005-02-16 株式会社東芝 タービンノズル
JP2001289003A (ja) * 2000-04-04 2001-10-19 Mitsubishi Heavy Ind Ltd ガスタービンの冷却構造
SE519353C2 (sv) * 2000-11-15 2003-02-18 Volvo Aero Corp Stator till en gasturbin
FR2825784B1 (fr) * 2001-06-06 2003-08-29 Snecma Moteurs Accrochage de chambre de combustion cmc de turbomachine utilisant les trous de dilution
US6554562B2 (en) * 2001-06-15 2003-04-29 Honeywell International, Inc. Combustor hot streak alignment for gas turbine engine
US6840048B2 (en) * 2002-09-26 2005-01-11 General Electric Company Dynamically uncoupled can combustor
JP3958728B2 (ja) * 2003-09-10 2007-08-15 株式会社日立製作所 ガスタービン燃焼器構造
GB0323909D0 (en) 2003-10-11 2003-11-12 Rolls Royce Plc Turbine blades
JP2005120871A (ja) 2003-10-15 2005-05-12 Mitsubishi Heavy Ind Ltd ガスタービン
US6929446B2 (en) 2003-10-22 2005-08-16 General Electric Company Counterbalanced flow turbine nozzle
JP4220947B2 (ja) * 2004-08-13 2009-02-04 三菱重工業株式会社 燃焼器尾筒とタービン入口との連通構造
JP4381276B2 (ja) 2004-10-08 2009-12-09 三菱重工業株式会社 ガスタービン
US7513102B2 (en) * 2005-06-06 2009-04-07 General Electric Company Integrated counterrotating turbofan
US7721547B2 (en) * 2005-06-27 2010-05-25 Siemens Energy, Inc. Combustion transition duct providing stage 1 tangential turning for turbine engines
US7377743B2 (en) * 2005-12-19 2008-05-27 General Electric Company Countercooled turbine nozzle
US7930891B1 (en) * 2007-05-10 2011-04-26 Florida Turbine Technologies, Inc. Transition duct with integral guide vanes
US7836703B2 (en) * 2007-06-20 2010-11-23 General Electric Company Reciprocal cooled turbine nozzle
US8387398B2 (en) * 2007-09-14 2013-03-05 Siemens Energy, Inc. Apparatus and method for controlling the secondary injection of fuel
JP2009197650A (ja) * 2008-02-20 2009-09-03 Mitsubishi Heavy Ind Ltd ガスタービン
US8065881B2 (en) * 2008-08-12 2011-11-29 Siemens Energy, Inc. Transition with a linear flow path with exhaust mouths for use in a gas turbine engine
US8033119B2 (en) * 2008-09-25 2011-10-11 Siemens Energy, Inc. Gas turbine transition duct
US7886991B2 (en) * 2008-10-03 2011-02-15 General Electric Company Premixed direct injection nozzle
US9822649B2 (en) * 2008-11-12 2017-11-21 General Electric Company Integrated combustor and stage 1 nozzle in a gas turbine and method
US8087253B2 (en) * 2008-11-20 2012-01-03 General Electric Company Methods, apparatus and systems concerning the circumferential clocking of turbine airfoils in relation to combustor cans and the flow of cooling air through the turbine hot gas flowpath
JP5180807B2 (ja) * 2008-12-24 2013-04-10 三菱重工業株式会社 １段静翼の冷却構造、及びガスタービン
DE102009011924A1 (de) * 2009-03-10 2010-09-16 Rolls-Royce Deutschland Ltd & Co Kg Nebenstromkanal eines Turbofantriebwerks
US8904799B2 (en) * 2009-05-25 2014-12-09 Majed Toqan Tangential combustor with vaneless turbine for use on gas turbine engines
US8082739B2 (en) * 2010-04-12 2011-12-27 General Electric Company Combustor exit temperature profile control via fuel staging and related method
US9267691B2 (en) * 2012-01-03 2016-02-23 General Electric Company Quick disconnect combustion endcover

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None *

Also Published As

Publication number Publication date
US9395085B2 (en) 2016-07-19
KR20130087640A (ko) 2013-08-06
WO2011070806A1 (ja) 2011-06-16
KR20120058600A (ko) 2012-06-07
KR101377772B1 (ko) 2014-03-25
JP2011117700A (ja) 2011-06-16
EP2511612A1 (en) 2012-10-17
JP5479058B2 (ja) 2014-04-23
KR101415478B1 (ko) 2014-07-04
CN102686949A (zh) 2012-09-19
EP2511612A4 (en) 2016-05-11
US20120247125A1 (en) 2012-10-04
CN102686949B (zh) 2015-02-04

Similar Documents

Publication Publication Date Title
EP2511612B1 (en) Gas turbine component assembly
US20240159151A1 (en) Airfoil for a turbine engine
US8523523B2 (en) Cooling arrangements
US8550774B2 (en) Turbine arrangement and method of cooling a shroud located at the tip of a turbine blade
US9011077B2 (en) Cooled airfoil in a turbine engine
US20180230836A1 (en) Stator vane section
EP2375161B1 (en) Combustor having a flow sleeve
KR101509385B1 (ko) 스월링 냉각 채널을 구비한 터빈 블레이드 및 그 냉각 방법
CN101220964A (zh) 套环、套筒以及冷却燃烧器组件的方法
EP2484872B1 (en) Passive cooling system for a turbomachine
EP2691610A1 (en) Turbine combustion system cooling scoop
JP2014009937A (ja) ガスタービン用移行ダクト
JP2003083001A (ja) ガスタービン及びその静翼
EP2955443B1 (en) Impingement cooled wall arrangement
US10378361B2 (en) Gas turbine blade
EP2859273B1 (en) Combustor liner with convergent cooling channel
JP2009013837A (ja) ガスタービン設備
US11655716B2 (en) Cooling structure for trailing edge of turbine blade
JP5614954B2 (ja) 燃焼器とタービン部との連通構造、および、ガスタービン
EP2771554B1 (en) Gas turbine and method for guiding compressed fluid in a gas turbine
US11585228B2 (en) Technique for cooling inner shroud of a gas turbine vane
JP2018150913A (ja) タービン翼及びそれを備えたガスタービン
JP2014005812A (ja) ガスタービン翼

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 20120314

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO SE SI SK SM TR

DAX Request for extension of the european patent (deleted)
RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: MITSUBISHI HITACHI POWER SYSTEMS, LTD.

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: MITSUBISHI HITACHI POWER SYSTEMS, LTD.

RA4 Supplementary search report drawn up and despatched (corrected)

Effective date: 20160412

RIC1 Information provided on ipc code assigned before grant

Ipc: F23R 3/42 20060101AFI20160406BHEP

Ipc: F02C 7/28 20060101ALI20160406BHEP

Ipc: F01D 9/06 20060101ALI20160406BHEP

Ipc: F01D 9/02 20060101ALI20160406BHEP

Ipc: F02C 7/18 20060101ALI20160406BHEP

Ipc: F23R 3/34 20060101ALI20160406BHEP

Ipc: F23R 3/60 20060101ALI20160406BHEP

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

RIC1 Information provided on ipc code assigned before grant

Ipc: F01D 9/02 20060101ALI20170222BHEP

Ipc: F02C 7/28 20060101ALI20170222BHEP

Ipc: F23R 3/60 20060101ALI20170222BHEP

Ipc: F23R 3/34 20060101ALI20170222BHEP

Ipc: F02C 7/18 20060101ALI20170222BHEP

Ipc: F23R 3/42 20060101AFI20170222BHEP

Ipc: F01D 9/06 20060101ALI20170222BHEP

INTG Intention to grant announced

Effective date: 20170322

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAJ Information related to disapproval of communication of intention to grant by the applicant or resumption of examination proceedings by the epo deleted

Free format text: ORIGINAL CODE: EPIDOSDIGR1

GRAL Information related to payment of fee for publishing/printing deleted

Free format text: ORIGINAL CODE: EPIDOSDIGR3

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

INTC Intention to grant announced (deleted)
INTG Intention to grant announced

Effective date: 20170816

GRAJ Information related to disapproval of communication of intention to grant by the applicant or resumption of examination proceedings by the epo deleted

Free format text: ORIGINAL CODE: EPIDOSDIGR1

GRAL Information related to payment of fee for publishing/printing deleted

Free format text: ORIGINAL CODE: EPIDOSDIGR3

GRAR Information related to intention to grant a patent recorded

Free format text: ORIGINAL CODE: EPIDOSNIGR71

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

INTC Intention to grant announced (deleted)
AK Designated contracting states

Kind code of ref document: B1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO SE SI SK SM TR

INTG Intention to grant announced

Effective date: 20171004

REG Reference to a national code

Ref country code: GB

Ref legal event code: FG4D

REG Reference to a national code

Ref country code: CH

Ref legal event code: EP

Ref country code: AT

Ref legal event code: REF

Ref document number: 944489

Country of ref document: AT

Kind code of ref document: T

Effective date: 20171115

REG Reference to a national code

Ref country code: IE

Ref legal event code: FG4D

REG Reference to a national code

Ref country code: DE

Ref legal event code: R096

Ref document number: 602010046622

Country of ref document: DE

REG Reference to a national code

Ref country code: NL

Ref legal event code: MP

Effective date: 20171108

REG Reference to a national code

Ref country code: LT

Ref legal event code: MG4D

REG Reference to a national code

Ref country code: AT

Ref legal event code: MK05

Ref document number: 944489

Country of ref document: AT

Kind code of ref document: T

Effective date: 20171108

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: ES

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: SE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: FI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: NO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20180208

Ref country code: LT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: NL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: HR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: IS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20180308

Ref country code: BG

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20180208

Ref country code: AT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: LV

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: GR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20180209

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: CZ

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: EE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: CY

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: DK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 602010046622

Country of ref document: DE

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: RO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: SM

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: IT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: PL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed

Effective date: 20180809

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

REG Reference to a national code

Ref country code: CH

Ref legal event code: PL

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 20180514

REG Reference to a national code

Ref country code: BE

Ref legal event code: MM

Effective date: 20180531

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MC

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

REG Reference to a national code

Ref country code: IE

Ref legal event code: MM4A

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CH

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180531

Ref country code: LI

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180531

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: LU

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180514

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180531

Ref country code: IE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180514

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180514

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: BE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180531

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20180514

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: TR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: PT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

Ref country code: HU

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT; INVALID AB INITIO

Effective date: 20100514

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MK

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20171108

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: AL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20171108

REG Reference to a national code

Ref country code: DE

Ref legal event code: R082

Ref document number: 602010046622

Country of ref document: DE

Representative=s name: HENKEL & PARTNER MBB PATENTANWALTSKANZLEI, REC, DE

Ref country code: DE

Ref legal event code: R081

Ref document number: 602010046622

Country of ref document: DE

Owner name: MITSUBISHI POWER, LTD., JP

Free format text: FORMER OWNER: MITSUBISHI HITACHI POWER SYSTEMS, LTD., YOKOHAMA, KANAGAWA, JP

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: DE

Payment date: 20230331

Year of fee payment: 14